Spacecraft and Control Method

ABSTRACT

A method of controlling a spacecraft is provided. The method includes one or more of calculating, by a control device, spacecraft position, attitude, and velocity, the spacecraft including a plurality of blade actuators controlling pitch for a plurality of blades, the plurality of blades extending radially away from a spacecraft core and including material configured to be deflected by solar pressure, receiving mission plan updates for the spacecraft, calculating an updated trajectory based on the position, attitude, velocity, mission plan updates, and past spacecraft behavior, generating maneuver parameters for the spacecraft from the updated trajectory, creating new blade pitch profiles for a plurality of blade actuators, from the maneuver parameters, sending controls corresponding to the new blade pitch profiles to the plurality of blade actuators, and transitioning from current blade pitch profiles to the new blade pitch profiles.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a Division of National Stage application Ser. No.16/309,149, filed Dec. 12, 2018, U.S. Pat. No. 11,338,942, under 35U.S.C. § 371, of International Patent Application No. PCT/US17/37543,filed on Jun. 14, 2017, which claims priority to earlier filedProvisional Application no. 62/349,941 filed Jun. 14, 2016 and entitled“HELIOGYRO DESIGN, OPERATION, AND APPLICATION”, the entire contents ofwhich are hereby incorporated by reference.

FIELD

The present invention is directed to apparatuses and control methods forspacecraft. In particular, the present invention is directed toapparatuses and methods for a Heliogyro and deployment/control thereof.

BACKGROUND

A spacecraft is a vehicle or machine designed to fly in space.Spacecraft are used for a variety of purposes, including communications,earth observation, meteorology, navigation, space colonization,planetary exploration, and transportation of humans and cargo. On asub-orbital spaceflight, a spacecraft enters space and then returns tothe Earth, without having gone into orbit. For orbital spaceflights,spacecraft enter closed orbits around the Earth or around othercelestial bodies. Spacecraft used for human spaceflight carry people onboard as crew or passengers from start or on orbit (space stations)only, whereas those used for robotic space missions operate eitherautonomously or telerobotically. Robotic spacecraft used to supportscientific research may be space probes. Robotic spacecraft that remainin orbit around a planetary body may be artificial satellites. Somespacecraft, such as Pioneer 10 and 11, Voyager 1 and 2, and NewHorizons, are on trajectories that leave the Earth's Solar System.Orbital spacecraft may be recoverable, or not. By method of reentry toEarth they may be divided in non-winged space capsules and wingedspaceplanes.

Unmanned spacecraft are spacecraft without people on board, used forunmanned spaceflight. Unmanned spacecraft may have varying levels ofautonomy from human input, they may be remote controlled, remote guidedor even autonomous, meaning they have a pre-programmed list ofoperations, which they will execute, unless otherwise instructed byEarth-based guidance. Autonomous spacecraft may also havedecision-making capabilities for determining the order of operationsperformed or the parameters governing operations. Many habitablespacecraft also have varying levels of robotic features. For example,the space stations Salyut 7 and Mir, and the ISS module Zarya werecapable of unmanned remote-guided station keeping, and docking maneuverswith both resupply craft and new modules. The most common unmannedspacecraft categories are robotic spacecraft, unmanned resupplyspacecraft, space probes, and space observatories. Not every unmannedspacecraft is a robotic spacecraft, for example a reflector ball is anon-robotic unmanned spacecraft.

SUMMARY

The present invention is directed to solving disadvantages of the priorart. In accordance with embodiments of the present invention, a methodof controlling a spacecraft is provided. The method includes one or moreof calculating, by a control device, spacecraft position, attitude, andvelocity, the spacecraft including a plurality of blade actuatorscontrolling pitch for a plurality of blades, the plurality of bladesextending radially away from a spacecraft core and including materialconfigured to be deflected by solar pressure, receiving mission planupdates for the spacecraft, calculating an updated trajectory based onthe position, attitude, velocity, mission plan updates, and pastspacecraft behavior, generating maneuver parameters for the spacecraftfrom the updated trajectory, creating new blade pitch profiles for aplurality of blade actuators, from the maneuver parameters, sendingcontrols corresponding to the new blade pitch profiles to the pluralityof blade actuators, and transitioning from current blade pitch profilesto the new blade pitch profiles.

One advantage of the present invention is that it provides a spacecraftwith a high thrust-to-weight ratio. A large blade surface area providesa large cumulative solar pressure as the thrust source. Centripetaltension and chord-wise battens provide stiffness instead of booms,resulting in a lighter system.

Another advantage of the present invention is that it allows completeattitude control. A Heliogyro utilizes spun sails or blades to providecontrol similar to a helicopter. Collective and cyclic pitch of theblades adjusts the spin vector and controls the thrust vector.

Another advantage of the present invention is that it provides forcompact storage of the spacecraft. Each blade individually rolls up on aspool for stowage, thereby packing into the most compact volumepossible. This type of stowage facilitates deployment and avoidsfolding, which can impart permanent wrinkles into the sail or bladematerial. Additionally, unrolling a blade is inherently more reliablethan unfolding and tensioning.

Another advantage of the present invention is it allows for thespacecraft to be un-deployed and moved to a different location withouthaving to move itself. That is, the deployed blades may be rolled uponto the blade deployers, the struts may be rolled up or folded, solarpanels may be folded, and the entire spacecraft may be packed into acompact storage space such as a CubeSat envelope for transport andre-deployment. This facilitates spacecraft re-use in space, perhapseliminating a need to deploy a different spacecraft somewhere else.

Another advantage of the present invention is that it provides forexcellent spacecraft scalability. Blade width, blade length and thenumber of blades provided by the Heliogyro can be scaled up. Dynamicsare consistent throughout the practical trade space.

Another advantage of the present invention is that it provides forexcellent ground testability. Blade assemblies mounted to a verticalspinning axis within a vacuum chamber allow for blade deployment andcentripetally stiffened blade pitch tests to verify blade models andhardware. This results in the blade subsystem reaching a NASA TechnologyReadiness Level (TRL) of 6 without leaving Earth. TRL is a method ofestimating technology maturity of Critical Technology Elements (CTE) ofa program during the technology acquisition process. They are determinedduring a Technology Readiness Assessment (TRA) that examines programconcepts, technology requirements, and demonstrated technologycapabilities. TRL are based on a scale from 1 to 9 with 9 being the mostmature technology. The use of TRLs enables consistent, uniformdiscussions of technical maturity across different types of technology.

Yet another advantage of the present invention is that it provides forredundancy and fault tolerance. Multiple blades allows for compensationin the case of an anomaly such as a mechanical or electrical failure ofa blade or damage to a blade from an impact.

Additional features and advantages of embodiments of the presentinvention will become more readily apparent from the followingdescription, particularly when taken together with the accompanyingdrawings. This overview is provided to introduce a selection of conceptsin a simplified form that are further described below in the DetailedDescription. It may be understood that this overview is not intended toidentify key features or essential features of the claimed subjectmatter, nor is it intended to be used to limit the scope of the claimedsubject matter.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a diagram illustrating main components of a strut Heliogyro inaccordance with embodiments of the present invention.

FIG. 2 is a block diagram illustrating main components of a ringHeliogyro in accordance with embodiments of the present invention.

FIG. 3 is a diagram illustrating Heliogyro spin using strut-mountedthrusters 300 in accordance with embodiments of the present invention.

FIG. 4 is a diagram illustrating Heliogyro blade deployer details inaccordance with embodiments of the present invention.

FIG. 5 is a diagram illustrating a Heliogyro spring-loaded attachmentpoint of view A-A in accordance with embodiments of the presentinvention.

FIG. 6 is a diagram illustrating a Heliogyro blade feedout system inaccordance with embodiments of the present invention.

FIG. 7 is a diagram illustrating a Heliogyro blade feedout system inaccordance with embodiments of the present invention.

FIG. 8 is a diagram illustrating a cross section of a Heliogyro bladeactuator in accordance with embodiments of the present invention.

FIG. 9 is a diagram illustrating a cross section of a Heliogyro axialbearing offload device in accordance with embodiments of the presentinvention

FIG. 10 is a diagram illustrating Heliogyro deployed blade details inaccordance with embodiments of the present invention.

FIG. 11 is a diagram illustrating Heliogyro blade clamp roller actuationsteps in accordance with embodiments of the present invention.

FIG. 12 is a diagram illustrating exemplary Heliogyro CubeSatconfigurations in accordance with embodiments of the present invention.

FIG. 13 is a diagram illustrating an exemplary Heliogyro CubeSatconfiguration after ejecting from a CubeSat canister in accordance withembodiments of the present invention.

FIG. 14 is a diagram illustrating an exemplary Heliogyro CubeSatconfiguration after deploying solar panels in accordance withembodiments of the present invention.

FIG. 15 is a diagram illustrating exemplary Heliogyro CubeSat strutdeployment in accordance with embodiments of the present invention.

FIG. 16 is a diagram illustrating an exemplary Heliogyro CubeSatconfiguration after fully extending struts in accordance withembodiments of the present invention.

FIG. 17 is a diagram illustrating an exemplary Heliogyro CubeSatconfiguration with aligned blade deployers in accordance withembodiments of the present invention.

FIG. 18 is a diagram illustrating an exemplary Heliogyro CubeSatconfiguration with fully deployed blades in accordance with embodimentsof the present invention.

FIG. 19 is a diagram illustrating an exemplary alternative HeliogyroCubeSat storage arrangement in accordance with embodiments of thepresent invention.

FIG. 20 is a diagram illustrating damper locations for an alternativeHeliogyro configuration in accordance with embodiments of the presentinvention.

FIG. 21 is a flowchart illustrating a process for deploying a Heliogyroin accordance with embodiments of the present invention.

FIG. 22A is a diagram illustrating a Heliogyro Cyclic Pitch Maneuver inaccordance with embodiments of the present invention.

FIG. 22B is a diagram illustrating a Heliogyro Collective-Cyclic PitchManeuver in accordance with embodiments of the present invention.

FIG. 22C is a diagram illustrating a Heliogyro Collective Pitch Maneuverin accordance with embodiments of the present invention.

FIG. 22D is a diagram illustrating a Heliogyro Half-P Pitch Maneuver inaccordance with embodiments of the present invention.

FIG. 23 is a block diagram illustrating a Heliogyro Control device inaccordance with embodiments of the present invention.

FIG. 24 is a block diagram illustrating primary function flow for aHeliogyro control system in accordance with embodiments of the presentinvention.

FIG. 25 is a block diagram illustrating Heliogyro control system modelcontrollers in accordance with embodiments of the present invention.

FIG. 26 is a diagram illustrating Heliogyro despin mechanism options inaccordance with embodiments of the present invention.

FIG. 27 is a diagram illustrating a side view of a Heliogyro despinmechanism in accordance with embodiments of the present invention.

DETAILED DESCRIPTION

Solar sails (also called light sails or photon sails) are a form ofspacecraft propulsion using radiation pressure exerted by sunlight onlarge mirrors. A useful analogy may be a sailing boat; the lightexerting a force on the mirrors is akin to a sail being blown by thewind. High-energy laser beams could be used as an alternative lightsource to exert much greater force than would be possible usingsunlight, a concept known as beam sailing. Solar sail craft offer thepossibility of low-cost operations combined with long operatinglifetimes. Since they have few moving parts and use no propellant, theycan potentially be used numerous times for back-to-back missionsdelivering payloads.

Solar sails use a phenomenon that has a proven, measured effect onspacecraft. Solar pressure affects all spacecraft, whether ininterplanetary space or in orbit around a planet or small body. Atypical spacecraft going to Mars, for example, may be displaced bythousands of kilometres by solar pressure over the course of itstrajectory, so the effects must be accounted for in trajectory planning.Solar pressure may also affect the orientation or attitude of aspacecraft. The total force exerted on an 800 by 800 meter solar sail,for example, is about 5 newtons (1.1 lb-f) at Earth's distance from thesun. This makes it a low-thrust propulsion system similar in some waysto spacecraft propelled by electric engines, but dissimilar as well, asit uses no propellant. The force is exerted constantly as long as solarradiation strikes the sail blades. The collective effect of the constantforce over time is great enough to be considered a viable manner ofpropelling a spacecraft.

A Heliogyro is a propulsion system that utilizes reflected solarpressure as its only means of propulsion and attitude control. Thecontinuous force generated by photons striking the windmill-like sail issufficient to boost the craft out of earth's gravity well, andeventually out of the solar system. Collective and cyclic pitch of sailblades vectors thrust and adjusts the spin axis as needed to providecomplete control authority. Although the pressure from the sun is lessthan the pressure from a piece of paper on an open hand, its persistentpush imparts an accumulated thrust greater than rocket propulsion canprovide. A Heliogyro is capable of interplanetary operation without theuse of propellant, extending a spacecraft's reach to places that arecurrently otherwise unattainable.

Referring now to FIG. 1, a diagram illustrating main components of astrut Heliogyro 100 in accordance with embodiments of the presentinvention is shown. A Heliogyro spacecraft may be configured in manydifferent arrangements. In a first arrangement, a plurality of blades108 extends radially outward from a central spacecraft core 104 or core,where the plurality of blades 108 are in a common plane and equallyradially spaced from each other. The spacecraft core 104 includes one ormore computing devices that deploy, un-deploy, and control actuatedsurfaces and operation of the Heliogyro 100. Spacecraft core 104 mayalso include power storage devices to store electrical power. In someembodiments, spacecraft core 104 may also include one or more payloads,which may include cameras, sensors, experiments, antenna, and so forth.

In some embodiments, the Heliogyro 100 may include one or more powersources 120. In one embodiment, power sources 120 include one or moresolar panels. In a preferred embodiment, power sources 120 include oneor more solar panels integrated into the blade 108 material. Powersources 120 may provide power to operate the Heliogyro 100 to chargestorage devices of the spacecraft core 104, or for transfer to anotherspacecraft or entity on a planetary body.

In the embodiment illustrated, the Heliogyro 100 includes struts orstrut members 112 extending radially outward from the spacecraft core104. Struts 112 provide standoff for each of the blades from thespacecraft core 104 so that each blade 108 may be deployed independentlyfrom each of the other blades 108, without interference. Struts alsoprovide an offset between the blade root and spacecraft core 104, whichallows for blade damping out of the plane of the nominal core-strutrotation, via strut 112 bending or actuation with respect to thespacecraft core 104 and blade 108. In some embodiments, the struts 112include pairs of struts 112 extending in opposite directions away fromthe spacecraft core 104. In some embodiments, the struts may beextendable structures including booms and Storable Tubular ExtensionMember (STEM) booms.

Each strut 112 terminates in a blade deployer 116, which includes pitchactuators 124 to deploy and un-deploy each blade 108 of the Heliogyro100. In an un-deployed state, each blade 108 may be stored in a rolledconfiguration on blade deployers 116. In a deployed state, each blade108 is unrolled and extends radially outward from each blade deployer116. That is, when un-deployed, a blade 108 may be unrolled and fullyextended as a deployed blade 108, and once deployed, a blade 108 may berolled on a spindle or blade spool roll 704 of a blade deployer 116 intoan un-deployed state. Details of blade deployers 116 are described inmore detail with reference to FIGS. 4-7. Exemplary descriptions of eachof the primary Heliogyro 100 elements are provided in the followingparagraphs.

A Heliogyro 100 has the general appearance of a windmill and employssail control akin to a helicopter. A number of solar reflecting blades108 extend radially in the same plane and attach to a central bus viaextendable struts 112. In some embodiments, material for blades 108 mayconvert solar energy into electrical energy and thereby serve as solarblades 108. In one embodiment applicable to a CubeSat mission size, eachdeployed blade 108 measures 152 meters long, is up to 0.28 meters wide,and is constructed from aluminized, flight qualified 3 μm Kaptonpolyimide film. During operation, centripetal tension and chord-wisebattens provide stiffness to the deployed blades 108. A Heliogyrocontrol system 2400 uses collective and cyclic pitch of the deployedblades 108 to control attitude and thrust, and an onboard supervisedautonomous Guidance, Navigation, and Control (GNC) system to generateshort-term trajectory-following command sequences. In some embodiments,voice coil actuators 408 pitch and actively dampen the deployed blade108 from the root. In one embodiment applicable to a CubeSat missionsize, the Heliogyro 100 propulsion system weighs approximately 2.5kilograms. When fully deployed, characteristic thrust, or the thrustachieved by the Heliogyro 100 with its sails/blades normal to the Sun at1 Astronomical Unit (AU), is 0.12 milliNewtons (mN).

Referring now to FIG. 2, a block diagram illustrating main components ofa ring Heliogyro 200 in accordance with embodiments of the presentinvention is shown. A ring Heliogyro 200 includes similar blades 108,pitch actuators 124, and blade deployers 116 shown and described withreference to the strut Heliogyro 100 of FIG. 1. In a ring Heliogyro 200,the spacecraft core 204 is constructed from a ring of segments connectedby joints 208 in a ring arrangement, where each segment is mated toother segments at each end by joints 208. The joints 208 may be eitherflexible joints or a set of joints (i.e. revolute, spherical, universal,etc). The joints 208 may include in some embodiments active damping toreduce in-plane and/or out-of-plane vibration, and incorporate shapechanging to reduce modal disturbances to reduce or dampen unwantedvibrations of the spacecraft core 104, 204.

In some embodiments, the spacecraft core 204 is joined to blades 108with struts 112. In other embodiments, the segments of the spacecraftcore 204 may include pitch actuators 124 and blade deployers 116, suchthat struts 112 may not be present.

In some embodiments, an actuated despin mechanism 212 may be includedwithin the spacecraft core 204 structure, such as joined to oppositesegments. The despin mechanism 212 may include an inner core, with innergimbal actuation configured like a voice coil actuator of blade actuator408. The despin mechanism 212 may include a race or two of bearingsresiding between an inner core and an inner gimbal. Each axis mayinclude common slip rings. In the preferred embodiment, the despinmechanism 212 is gimbaled. In other embodiments, the despin mechanism212 is non-gimbaled and generally spins in just one direction. In someembodiments, the spacecraft core 204 wirelessly transmits power andcontrol data to the inner gimbal to control the actuated despinmechanism 212 and to communicate back and forth between the core andavionics or payloads on the despun section.

Referring now to FIG. 3, a diagram illustrating Heliogyro spin usingstrut-mounted thrusters 300 in accordance with embodiments of thepresent invention is shown. Heliogyro 100, 200 may include smallthrusters as part of propulsion modules 304. Although one propulsionmodule 304 may be able to spin up the he Heliogyro 100, 200 to apreferred spin rate, it may be more efficient to use a pair ofpropulsion modules 304A and 304B located 180 degrees apart either onopposite struts 112 or opposite blade deployers 116 located on oppositesides of the spacecraft core 104, 204. For example, of a six bladeHeliogyro 100, 200, spin up propulsion modules 304 should preferably beplaced on blades extending in opposite directions.

In the embodiment illustrated in FIG. 3, propulsion modules 304 arelocated in proximity to the outer ends of struts 112. In the embodimentillustrated in FIG. 4, propulsion modules 304 are located in proximityto the blade deployers 116. By positioning the propulsion modules 304 inproximity to the ends of the struts 112 or the blade deployers 116,thrust from the propulsion modules 304 will have a greater moment aroundthe spacecraft core 104, and allow smaller thrusters to be used then ifplaced closer to the spacecraft core 104, 204 itself In the preferredembodiment, each propulsion module 304 has a set of thrusters pointed indifferent directions (not shown for clarity) that, when turned onsingly, in pairs, or in some other pattern, can supply a net thrust in arange of directions in order to detumble the Heliogyro 100, 200 andorient it toward the sun.

Propulsion modules 304 provide a limited amount of thrust in a thrustdirection 308 to facilitate rotation of the Heliogyro 100, 200 aroundthe spacecraft core 104, 204. Although FIG. 3 illustrates a Heliogyro100, 200 rotating in a counterclockwise direction, it should beunderstood that the propulsion modules 304 may be configured in order tospin the Heliogyro 100, 200 clockwise instead. In the preferredembodiment, spin-up of the Heliogyro 100 200 is performed prior to theblades being deployed 316. Spin-up prior to blades 108 being deployedsupplies the necessary outward force, due to centripetal force, to pullthe blades 108 radially outward as each blade 108 is unrolled. Theinitial spin rate provides sufficient pull for a significant portion ofthe blade deployment. Once spun-up, the Heliogyro 100, 200 spins in aspin direction 312 within the plane of the Heliogyro 100, 200.

As an alternative to propulsion modules 304, a rocket upper stage mayhave a spinning apparatus that pre-spins a Heliogyro 100, 200 as it isdeployed. As another alternative, reaction wheels or magneticapparatuses may be used in lieu of thrusters. In addition to providingan initial spin-up of the Heliogyro 100, 200, a propulsion module 304can orient the Heliogyro 100, 200 normal to the sun, where it can beefficiently utilized once fully deployed.

Referring now to FIG. 4, a diagram illustrating Heliogyro blade deployer116 details in accordance with embodiments of the present invention isshown. Heliogyro blade deployers 116 provide storage for un-deployedblades 108 in rolled form and control apparatuses to deploy andun-deploy each of the blades 108. A blade roll 416 stores the entirelength of the blade 108 when the blade 108 is not deployed. The bladeroll 416 provides the most efficient form for storage of an un-deployedblade 108 compared to folding or articulating a long and thin structuresuch as a blade 108. The blade roll 416 is retained between a pair ofend caps 420. The end caps 420 capture each end of a spindle or bladeroll spool 704 about which the blade roll 416 is stored. The end caps420 are joined by a rigid structure extending across the width of theblade roll 416, which is generally joined to the pitch actuator 124,which attaches to the end of a strut 112 for strut Heliogyros 100 orring Heliogyros 200 using struts 112. Each blade deployer 116 may alsoinclude a propulsion module 304, as previously discussed with respect toFIG. 3. View A-A is provided in FIG. 5 to show additional detail of thepitch actuator 124 mounting.

Pitch for each of the blades 108 is controlled by a pitch actuator 124coupled to the widthwise support structure of the blade deployer 116. Inthe preferred embodiment, the pitch actuator 124 is a Lorentz Coil,which is a wire coil around a magnet - similar to a brushed motorwithout commutation. An audio voice coil motor is an example of this,which is preferred for less friction and higher lifetime (many millionsof cycles). In less-desirable actuators such as motors utilizing brushesor ball bearings, brushes and ball bearings wear too fast. Also,piezoelectric transducers may be used, but this disadvantageouslyrequires more power than Lorentz coil devices. Another advantage tousing Lorentz coil actuators in conjunction with the blade deployers 116is they may provide resistance or electromagnetic force as feedback thattells the control system how the blade 108 is performing.

In order to control the orderly deploying and un-deploying of the blade108 material, a blade clamp roller 412 or similar structure is providedon the blade deployer 116. In one embodiment, a blade clamp roller 412is always in contact with the blade roll 416, and exerts constantpressure to the blade roll 416 under spring force (not shown). Inanother embodiment, the blade clamp roller 412 has an associated bladeclamp roller actuator (not shown) that may either clamp to the bladeroll 416 or un-clamp and not be in contact with the blade roll 416. Thisactuated embodiment may be preferable by allowing the blade roll 416 todeploy the blade 108 under centripetal force when the blade clamp roller412 is unclamped, as the Heliogyro 100, 200 is rotating or turning. Atany point during the blade 108 deployment, the blade clamp roller 412may be actuated to engage the blade roll 416 and limit or preventfurther deployment of the blade 108. One alternative to the blade clamproller 412 is a blade clamp bar (not shown), which positively clamps thedeployed blade 108 to the blade roll 416 to fix the blade 108 at acurrent length of deployment from the blade deployer 116.

Referring now to FIG. 5, a diagram illustrating a Heliogyrospring-loaded attachment point of view A-A in accordance withembodiments of the present invention is shown. FIG. 5 illustrates a rearview of a blade deployer 116, from the view perspective of thespacecraft core 104, 204. The blade deployer 116 includes a bladeassembly yoke 512, which spans the width of the blade deployer 116 andprovides structure to support the blade roll 416 and blade clamp roller412. In the preferred embodiment, each blade deployer 116 includes apitch actuator 124 to rotationally clockwise or counterclockwise pitch520 a deployed blade 108 relative to a pitch actuator 124. Also shown inFIG. 5 is a cross-section of the pitch actuator yoke 504.

It is important to be able to actively or passively damp any unwantedfrequencies and motion in all three axes. When the spacecraft 104, 204is operating, the in-plane motion (i.e. within the spin plane) and theout-of-plane motion (i.e. perpendicular to the spin plane) need to bedamped. An in-plane damping system may be provided in the illustrateddirection 508 where the strut 112 meets the blade deployer 116, forexample using a sliding spring-loaded attachment point 516 as shown thatmoves horizontally within an attachment point slot 524. Alternatively,various motors or actuators could be used to provide in-plane damping.The spring-loaded attachment point 516 allows for two things: first, itallows the blade deployer 116 to shift sideways to aid in fitting insidea CubeSat or other stowage envelope; second, it can be embodied as aspring-damper, and can thus dampen out in-plane blade vibration modes508.

Referring now to FIG. 6, a diagram illustrating a blade deployer 116feedout system in accordance with embodiments of the present inventionis shown. The blade deployer 116 includes a blade roll 416 and a bladeclamp roller 412, and attaches to a distal end of a strut 112. The strut112 is a rigid structure providing standoff 604 from the spacecraft core104, 204 to the blade deployer 116. Struts 112 provide an improvementover conventional Heliogyro proposals that utilize combinations of polesand wires that extend from the blade assembly yoke 512 outward to alocation along the blade length to add rigidity and dampen out-of-planemovement. Struts 112 may in some embodiments advantageously require lessstorage area than poles/wires, and snags during blade 108 deployment areless likely. In the preferred embodiment, the strut 112 is between 1% to3% of the length of the deployed blade 108. Therefore, for a deployedblade 108 1000 meters in length, the strut 112 would normally beexpected to be between 10 meters to 30 meters in length.

The strut 112 may be a STEM boom or an unrolled section of rigidizablematerial that is stored in a compact fashion such as a roll when in anun-deployed state. In some embodiments, a blade hinge is offset from thecore to dampen both in-plane and out-of-plane vibrations. The bladehinge is the point on the blade roll 416 where the deployed blade 108departs. The blade 108 is so thin that the departure point acts like ahinge. The blade 108 can easily bend up and down at that point. Thedistance between the blade hinge and the strut attachment point on thespacecraft core 104, 204 is the blade hinge offset 604, or in otherwords, the offset distance from the spacecraft core 104, 204.

The strut 112 replaces other structures of the conventional artinvolving various forms of wires to provide stiffness and out-of-planedamping. Either in-plane and/or out-of-plane dampers may be located atthe root of each strut 112, where it attaches to the spacecraft core104, 204.

Referring now to FIG. 7, a diagram illustrating a Heliogyro blade feedout system in accordance with embodiments of the present invention isshown. FIG. 7 illustrates a blade deployer 116 without the blade 108 orblade roll 416 shown, in order to represent the supporting componentsmore clearly. The blade roll 416 is stored on a blade roll spool 704 orspindle, which allows the blade roll 416 to be stored in the mostefficient space possible. Rotation of the blade roll spool 704 iscontrolled by a blade deployment actuator 708 within one or both endcaps 420 or the blade roll spool 704. The blade deployment actuator 708provides a lightweight and compact arrangement to unroll the blade 108at a carefully controlled rate. In some embodiments, the bladedeployment actuator 708 is within the blade roll spool 704 and drivesthe spool via an engagement feature, half of which resides on theactuator shaft and the other half is attached to the blade roll spool704.

The blade deployment actuator 708 may be any form of suitable radialactuator including a DC motor, a stepper motor, or similar. Onealternative to a blade deployment actuator 708 is to provide a mass atthe end of each blade 108, which will auto-deploy the blade 108 undercentripetal or Coriolis force when the blade 108 is spinning. In oneembodiment, the blade deployer 116 may know when the blade 108 is fullydeployed or half deployed or at any other position is by providing anoptical code such as a barcode or other known optical code on the blade108 material itself, and a corresponding optical sensor within the bladedeployer 116 apparatus. When an appropriate portion of blade 108material passes under the optical sensor, the optical code will bedetected by the sensor, which may then adjust the blade clamp roller 412or the blade deployment actuator 708, as needed. As an alternative, anencoder or other form of potentiometer in proximity to blade deploymentactuator 708 or blade roll spool 704 may be used instead.

Referring now to FIG. 8, a diagram illustrating a Heliogyro pitchactuator 408 in accordance with embodiments of the present invention isshown. A pitch actuator 124 is located at the root of a blade 108 (i,e,where a blade deployer 116 is attached to a strut 112), at the center ofa blade deployer 116. In the preferred embodiment, blade pitch dampingis provided by a voice coil rotary actuator or Lorenz coil with anencoder 804.

Voice coil actuators, or Lorentz coils, advantageously provide rapidresponse, low friction, and high reliability. The pitch actuator 124provides blade pitch and active torsional damping. Since no gearing isrequired by a voice coil actuator, the inherently low friction allowsmechanical feedback from the blade 108 to the pitch actuator 124.

Each Heliogyro pitch actuator 124 includes a pitch actuator shaft 808which rotates about a central axis through the pitch actuator 124device. In the voice coil configuration, pitch actuator control signals2344 energize a coil 820 and magnet 824 coupled to the pitch actuatorshaft 808 in order to rotate the shaft 808 a predetermined amount ineither direction. The pitch actuator 124 includes a rear bearing 812 anda front bearing 816, which are preferably very low friction bearings.Pitch actuator 124 may also include an encoder 828 to provide feedbackas to a rotary position of the pitch actuator shaft 808.

Referring now to FIG. 9, a diagram illustrating a Heliogyro axialbearing offload device in accordance with embodiments of the presentinvention is shown. As previously described, the pitch actuator 124 maybe a Lorentz coil or similar type of voice coil device, and may includean encoder 828 or other type of rotary position sensor. FIG. 9illustrates a special case of the pitch actuator 124 shown and describedwith reference to FIG. 8.

In order to minimize axial forces on the front bearing 916 and rearbearing 912, a retainer 932 and thread 928 provide inward radial forceopposite to outward centripetal force applied to the pitch actuatorshaft 908. The thread 928 provides a small amount of axial force to theshaft of the actuator in order to minimize friction between the shaft908 and the two bearings 912, 916 shown. The thread 928 is preferablymade from steel, Kevlar, or other material with high axial strength andminimal torsional resistance. Centripetal force of the spinningHeliogyro 100, 200 causes the shaft 904 to exert force outward (to theright, as shown). The thread 928 is intended to provide an equal andopposite force to the shaft 908, thus minimizing frictional forcesincluding torsional resistance to the bearings 912, 916, improvingtorque feedback from the blade 108, and increasing bearing life.

Referring now to FIG. 10, a diagram illustrating Heliogyro deployedblade 108 details in accordance with embodiments of the presentinvention is shown. In general, a Heliogyro deployed blade 108 includesa long strip of material that is both flexible enough to be stored on ablade roll spool 704 of a blade deployer 116 and yet rigid enough, dueto centripetal stiffness, to maintain shape and orientation whenrotated. The blade material is so thin that it has essentially zerostiffness. The only “effective stiffness” is therefore due tocentripetal force.

In the preferred embodiment, deployed blades 108 include one or moreaccelerometers at the distal or far end of the blade 108. In a preferredembodiment, two accelerometers are within a blade batten 1008 or otherstiff material across the width of the blade tip 1016. Differentialacceleration data from two accelerometers help to determine blade 108twist at the blade tip 1016. Accelerometers in the blade tip 1016 eithersends signals back to a blade deployer 116 or the spacecraft core 104,204 through an RF signal or deposited wiring on the surface 1012 of thedeployed blade 108. Accelerometers may provide information to theHeliogyro control system 2400 of what the blade tip 1016 is doingrelative to the root of the blade 108, where the blade 108 joins theblade deployer 116. In one embodiment, the accelerometers may begyroscopes or similar devices. In some embodiments, amplifiers spacedalong a blade edge may be required to boost the accelerometer signalsfor accurate reception by the blade deployer 116 or spacecraft core 104,204. In some embodiments, a miniature transmitter may be placed in closeproximity to the accelerometers, such as a miniature radio transmitterco-located with the solar cells 1004, to transfer the required signalsas previously described.

In order to provide power to the accelerometer, a power source isrequired. In one embodiment, solar cells 1004 are provided in proximityto the blade tip 1016. In another embodiment, solar cells 1004 may beprovided at the blade root where the blade attaches to a blade deployer116. However, for solar cells 1004 installed at a blade root, therewould undoubtedly be resistive effects in wiring routed from the solarcells at the root to sensors at the tip, increasing the required voltageto power the accelerometers at the blade tip 1016.

In a preferred embodiment, blade edge stiffeners 1020 may be provided onone or both lengthwise edges of blades 108 and may be bonded directly tothe blade 108 itself using Kevlar, polyimide, or similar material. Insome embodiments, thin polyimide material with a residual stressdifferential may be fused to one side of the blade 108 material.Residual stress differential is a difference in stress between the frontand back surfaces of blades 108, due to some chemical or physicalprocess, that causes the material to roll up into a small, tight tube.One edge is permanently fused and the other has a temporary glue thatweakens with time. When the weakened blade edge releases, the bladematerial rolls up into a cylindrical tube thus creating a stiffenededge. In other embodiments, blade edge stiffeners 1020 may be lengthwisetubes or STEM boom structures attached to the lengthwise blade edges. Asthe blades 108 deploy, the STEM boom rolls up into cylinders. Blade edgestiffeners 1020 may be uniform structures along the entire length of adeployed blade 108, or many sections of blade edge stiffeners 1020.Blade tip 1016 includes a blade batten 1008, which is a tubularstructure across the width of the blade tip 1016 and attached to theblade tip 1016. The blade batten houses one or more accelerometers, andserves to resist tension/compression deformation due to the Poissoneffect.

In the embodiment of many sections of blade edge stiffeners 1020, bladebattens 1008 may be provided to transversely couple blade edgestiffeners 1020. In this way, the blade edge stiffeners 1020 and bladebattens 1008 would form a “ladder” of structural support along thelength of each deployed blade 108. Blade edge stiffeners 1020 and bladebattens 1008 may be made of the same type of material, or differentmaterials. Another benefit of the blade edge stiffeners 1020 is they mayact as a rip-stop material to prevent cuts or punctures from severingthe deployed blade 108.

Referring now to FIG. 11, a diagram illustrating Heliogyro blade clamproller actuation steps 1100 in accordance with embodiments of thepresent invention is shown. The illustrated actuation steps proceed inchronological order from (A) through (E) during blade 108 deployment. Ifa blade 108 is being un-deployed or stowed on the blade deployer 116,the blade clamp roller 412 stays in position E and the blade deploymentactuator 708 pulls the blade material between the blade clamp roller 412and the blade roll spool 704, thus keeping the blade 108 taut as itwraps around the blade roll spool 704.

Step (A) illustrates the relationship between the blade roll 416 and theblade clamp roller 412 prior to the blade being deployed 1116. The bladeclamp roller 412 is in a far rear position (i.e. oriented toward thespacecraft core 104, 204, and is not engaged with the blade roll 416. Atthis point, the entire blade 108 is stored on the blade roll 416.

The sequence of steps (B) through (E) illustrate the relationshipbetween the blade clamp roller 412 and the blade roll 416 after theblade 108 has been either partially or fully deployed to a maximumlength 1120. A spring-loaded arm 1112 is rotationally coupled underspring force from a wire spring 1104 to a roller support tab 1128 thatengages the blade roll spool 704 (see FIG. 7). The wire spring 1104drives the spring-loaded arm 1112 toward the blade roll spool 704.Before blade clamp roller 412 actuation, an actuator, mounted on theroller support tab 1128, engages the blade roll spool 704 andmechanically couples the roller support tab 1128 with the blade rollspool 704. As the blade deployment actuator 708 drives the blade rollspool 704 in the deploy direction, the roller support tab 1128 rotatesas well, moving the spring-loaded arm 1112 and the blade clamp roller412 “outward”. The spring-loaded arm 1112 rides along a curved ramp 1108that is mechanically coupled to the end caps 420, and is angled so thatthe spring-loaded arm 1112 causes the blade clamp roller 412 to beclamped to the blade roll spool 1124 at the end of blade clamp roller412 actuation. In the middle of blade 108 deployment, the blade clamproller 412 clamps to the blade 108 so it can be pitched, thus providingextra momentum as the blade 108 feeds out, which is used to keep theHeliogyro 100, 200 spin rate constant. In the preferred embodiment, theblade clamp roller 412 may also be disengaged from the blade roll 416 atany time, if desired. Alternately, a blade clamp roller 412 or bar canbe rotated into position via an actuator-driven hinge point locatedelsewhere on the blade assembly yoke 512 or end caps 420. Alternately,an actuator may be included within the Roller Support Tab 1128 to selectengagement/disengagement as required.

The blade clamp roller 412 is initially clamped to the blade roll 416,and is locked into place. During blade 108 deployment, the spring-loadedarm 1112 unlocks and backdrives a small amount to lift off the bladeroll 416. After the blade 108 has unrolled, either partially orcompletely, and the blade clamp roller 412 needs to clamp the root ofthe blade 108, a linear actuator mounted on the roller support tab 1128engages the blade roll spool 704, thus driving the spring-loaded arm1112 forward. As it moves, the spring-loaded arm 1112 rotates inward,riding along the curved ramp 1108, and eventually resting the bladeclamp roller 412 on the blade roll 416. Clamping holds blade 108material at the root in place to allow for pitching of the blade 108. Italso allows for reversing blade roll 416 motion in order to re-stow theblade 108.

The Heliogyro 100, 200 initially spins up to a spin rate that providessufficient centripetal force to pull a first portion of the blades 108out and keep the blades 108 radially extended. As the blades 108 feedout, the spin rate slows. At a certain point, when the spin rate isnearing a level that won't support controlled radial deployment, theblade clamp rollers 412 clamp the blades 108 and the blades 108 arepitched. Solar pressure is impacting the blades 108 constantly up tothis point, but now, with the blades 108 pitched, the solar pressureadds angular momentum. The blades 108 can then be fed out at a rate thatmaintains the necessary spin rate, with margin, to ensure a controlled,radial deployment.

Referring now to FIG. 12, a diagram illustrating exemplary Heliogyro100, 200 CubeSat configurations in accordance with embodiments of thepresent invention is shown. The illustrated configurations reflect aHeliogyro 100, 200 with four blade rolls 416 and blade deployers 116.For stowage, each blade 108 is rolled onto a blade roll spool 704adjacent to its pitch actuator 408. The blade roll assemblies 416 andstruts 112 fold up in three ways, depending on the CubeSat volumeallocated: (A) along the bottom 2U of a 6U volume, (B) along oppositesides of the spacecraft core 104, 204, or (C) stacked on its end.

Referring now to FIG. 13, a diagram illustrating an exemplary Heliogyro100, 200 CubeSat configuration after ejecting from a CubeSat canister inaccordance with embodiments of the present invention is shown. AHeliogyro 100, 200 in CubeSat format must be ejected from a CubeSatcanister of a space vehicle before the CubeSat can be deployed. FIG. 13illustrates an example of an external appearance of a CubeSat Heliogyro100, 200 prior to deployment. One side of the spacecraft core 104 isvisible, with the bulk of the spacecraft core 104, 204 verticallyoriented in the center of the CubeSat. The large sides of the CubeSatare solar panels of the power source 120, and the blade rolls 416 andblade deployers 116 are on the small sides of the CubeSat. In theembodiment illustrated, there are six blade deployers 116 and bladerolls 416.

Referring now to FIG. 14, a diagram illustrating an exemplary Heliogyro100, 200 CubeSat configuration after deploying solar panels 120 inaccordance with embodiments of the present invention is shown. The powersource 120, shown as solar panels, are hinged at the top surfaces in a“gull wing” configuration and attached to the spacecraft core 104, 204.This frees up space for the blade deployers 116 and blade rolls 416 tobe extended by struts 112 (not shown), prior to the blades 108 beingdeployed. Once the power source 120 is deployed, the solar panels are ina planar configuration.

Referring now to FIG. 15, a diagram illustrating exemplary Heliogyro 100CubeSat strut 112 deployment in accordance with embodiments of thepresent invention is shown. In the illustrated embodiment, each of thestruts 112 is composed of a number of hinged segments which are storedin a folded configuration on the CubeSat. In one embodiment, each of thestruts 112 is stored in a roll, and are unrolled to a fully extendedlength. In another embodiment, each of the struts 112 are STEM booms. Inanother embodiment, each of the struts 112 is longitudinally collapsedfor stowage, and extend outward for deployment. In one embodiment, eachof the strut 112 segments are spring-actuated. Once the power source 120has been deployed and is out of the way, the spacecraft core 104 rotateseach of the struts 112 to provide equal spacing between each of thestruts 112.

Referring now to FIG. 16, a diagram illustrating an exemplary Heliogyro100, 200 CubeSat configuration after fully extending struts 112 inaccordance with embodiments of the present invention is shown. Each ofthe struts 112 are extended to a maximum length. The length ispredetermined, and is generally 1 to 3% of the total blade 108 length.At this step, propulsion units on the struts 112, such as propulsionmodules 304, detumble the Heliogyro 100, 200 and point it toward thesun.

Referring now to FIG. 17, a diagram illustrating an exemplary Heliogyro100, 200 CubeSat configuration with aligned blade deployers 116 inaccordance with embodiments of the present invention is shown. Thespacecraft core 104 rotates each of the blade deployers 116 90° suchthat deployed blades 108 will be normal to the sun. While maintaining alock on the sun, the propulsion units 304 spin up the Heliogyro 100, 200to a first spin rate, preferably 60 revolutions per minute. A practicalminimum spin rate ensures sufficient centripetal force to pull theblades 108 outward, off of the blade roll 416. A practical maximum spinrate keeps the stress in the blades 108 below safe limits. These are afunction of the weight of the blade tip batten assembly 1008 (the weightat the blade tip 1016). 60 rpm is sufficient to deploy the blades 108 asignificant distance of the way before having to pitch the blades 108and allow solar pressure to boost angular momentum. This advantageouslycuts down on required deployment time.

Referring now to FIG. 18, a diagram illustrating an exemplary Heliogyro100, 200 CubeSat configuration with fully deployed blades 108 inaccordance with embodiments of the present invention is shown. Once theblade deployers 116 have been aligned, the blades 108 feed out in acontrolled, balanced manner. The feed rate is sufficiently slow tomaintain a positive trailing edge stress in the blades 108, thusguaranteeing a smooth, radial deployment. If the blades 108 feed out toofast, they bend backward with respect to the spin direction. This isbecause the trailing edge no longer has tension, but is placed intocompression. Coriolis forces cause the blades 108 to lag behind therotation of the Heliogyro 100, 200. This is countered by radialcentripetal stress, but the spin rate is limited to a certain amount.

During deployment, the Heliogyro 100, 200 slows to its operating spinrate, or second spin rate, at which time the blades 108 collectivelypitch as directed by the control system 2400. The spin rate may be onthe order of 1 rpm, however, for blades 108 less than 1000 meters long.The operating or second spin rate is chosen to limit blade 108 coning(droop out of plane) to a certain amount, which in turn limits thecoupling between blade twist, in-plane deflection, and out-of-planedeflection. It is also chosen to still allow adequate precession ratesto aid maneuverability.

For the rest of the deployment, solar pressure on the blades 108 booststhe angular momentum while the blades 108 feed out at a speed necessaryto maintain a constant spin rate. As the blades 108 feed out, if solarpressure doesn't act to boost it, the angular momentum is constant. Thismeans the spin rate slows as the blades 108 deploy. The change in spinrate is calculated by taking the initial angular momentum, with theblades 108 at a certain length, and then setting the momentum equal tothe equation for angular momentum if the blades 108 are longer. Theequation will show that, if moment of inertia is larger (as is the casewhen the blades 108 are extended out farther), then the spin rate isless. Knowing how the spin rate varies, allows one to determine how muchthe blades 108 need to be pitched to support a certain blade feed-outrate.

Referring now to FIG. 19, a diagram illustrating an exemplaryalternative Heliogyro 100, 200 CubeSat storage arrangement in accordancewith embodiments of the present invention is shown. The Heliogyro 100,200 configuration illustrated in FIG. 19 shows four blade deployers 116,with each blade deployer 116 including a blade clamp roller 412 and apitch actuator 408.

Referring now to FIG. 20, a diagram illustrating damper locations for analternative Heliogyro 100, 200 configuration in accordance withembodiments of the present invention is shown. FIG. 20 shows astructural support 2008 and four in-plane and out-of-plane dampers 2004,as well as rolled up struts 112 for each of the blade deployers 116 andblades 108. The in-plane and out-of-plane damper assemblies 2004 areeither passive or active. The passive damper configurations have springsand damper materials, either in series or parallel, that act to dampenunwanted motion. They can also have coils of wires and magnets thatresist and dampen motion due to electromotive force generated as amagnet moves with respect to a wire. Active damper configurations aresimilar to voice coils or some other motor configuration. Whenback-driven, or driven against the powered direction of the motor oractuator, they encounter resistance. By tuning the resistance of thecoil, motor, and voltage applied to the active damper motor circuit, theamount of damping is tuned to an optimal level.

Referring now to FIG. 21, a flowchart illustrating a process 2100 fordeploying a Heliogyro in accordance with embodiments of the presentinvention is shown. Flow begins at block 2104.

At block 2104, a space vehicle ejects the Heliogyro 100, 200. In oneembodiment, the Heliogyro 100, 200 is stored as a CubeSat. In anotherembodiment, the Heliogyro 100, 200 is stored as a non-CubeSat payload ona space vehicle. Flow continues to block 2108.

At block 2108, propulsion modules 304 detumble the Heliogyro 100, 200and orient the Heliogyro 100, 200 toward the sun. The propulsion modules304 are either mounted on struts 112 or blade deployers 116. TheHeliogyro 100, 200 is oriented toward the sun when the blades 108, whendeployed, are normal to the sun. Flow proceeds to block 2112 and block2116.

At optional block 2112, for Heliogyros 100, 200 that include struts 112,each of the struts 112 are deployed and extended. When the struts 112are deployed or extended (for example, per FIGS. 15 and 16), the struts112 are unrolled, extended, or unfolded into a linear disposition. Someconfigurations of Heliogyro 100, 200, such as a ring configurationHeliogyro 200 that incorporates blade deployers 116 as the spacecraftcore 204, may not include struts 112. Deployment may include rotatingeach of the struts 112 to maintain equal angular spacing between thestruts 112. Flow proceeds to block 2116.

At block 2116, the Heliogyro control system 2400 spins the Heliogyro100, 200 to a first spin rate, which in the preferred embodiment isapproximately 60 RPM. The Heliogyro 100, 200 spins within a common planegenerally defined by the spacecraft core 104, 204, struts 112, bladedeployers 116, and blades 108. Flow proceeds to block 2120.

At block 2120, the Heliogyro control system 2400 begins to feed out eachof the blades 108. As previously discussed, blades 108 are fed out at apredetermined rate in order to maintain a trailing edge stress on eachof the blades 108. Flow proceeds to decision block 2124.

At decision block 2124, as the blades 108 are extending, the Heliogyrocontrol system 2400 determines if the blade spin rate has achieved asecond spin rate. In the preferred embodiment, the second spin rate isapproximately one RPM. If the blade spin rate has not achieved thesecond spin rate, then flow proceeds to decision block 2124 to keepchecking the spin rate. If the blade spin rate has achieved the secondspin rate, then flow instead proceeds to block 2128.

At block 2128, the Heliogyro control system 2400 pitches the blades 108.In one embodiment, each of the blades 108 is pitched to a predeterminedpitch angle. Flow proceeds to block 2132.

At block 2132, the Heliogyro control system 2400 continues feeding outthe blades 108 to a fully extended length. At this point, the Heliogyro100, 200 is fully deployed and operational. Flow ends at block 2132.

Referring now to FIG. 22A, a diagram illustrating a Heliogyro CyclicPitch Maneuver in accordance with embodiments of the present inventionis shown. FIG. 22A illustrates cyclic pitch, which results in in-planethrust in a specific direction 2216. FIGS. 22A-22D also includes a PitchAngle vs Time graph 2204, 2220, 2232, 2236 for the deployed blades 108,and assume a Heliogyro 100, 200 with thrust and spin direction as shownin FIG. 3. The horizontal line in the graph represents a pitch angle of0°, where the blades 108 are unpitched. A dot at the Y-axis locationindicates the pitch angle for the darkened (black) blade. All bladesfollow the same profile, albeit at different places on the profile. Forexample, the dark blade is at phase 0°. The next blade goingcounterclockwise is 30° ahead in phase. Its pitch angle is defined bythe location on the pitch profile 1/12 of a period ahead in time (30°divided by 360°). The next blade proceeding counterclockwise is 60°ahead of the first blade and 30° ahead of the second blade.

The uniform motion of each blade 108 is due to its fundamental pitchfrequency of approximately one cycle per Heliogyro 100, 200 revolution,which is the basic frequency at which the blades 108 are pitched.Amplitude of pitch is typically on the order of 10°, and ranges from 0°to 90°.

Referring now to FIG. 22B, a diagram illustrating a HeliogyroCollective-Cyclic Pitch Maneuver in accordance with embodiments of thepresent invention is shown. FIG. 22B illustrates collective-cyclicpitch, which changes the Heliogyro 100, 200 spin rate and precesses aspin vector for the Heliogyro 100, 200. The Heliogyro 100, 200precesses, or rotates, its spin vector due to torque imparted by the sunimpacting the blades 108 as they pitch along the pitch angle vs timeprofile 2220. Note that the collective, or average pitch angle vs timeis offset from 0°. This offset can be either positive (above the line)or negative (below the line).

Referring now to FIG. 22C, a diagram illustrating a Heliogyro CollectivePitch Maneuver in accordance with embodiments of the present inventionis shown. FIG. 22C illustrates collective pitch, which changes theHeliogyro 100, 200 spin rate 2224. Note that the pitch angle vs time isconstant and offset from 0°—meaning the blades 108 are pitched and thepitch does not change during the maneuver.

Referring now to FIG. 22D, a diagram illustrating a Heliogyro Half-PPitch Maneuver in accordance with embodiments of the present inventionis shown. FIG. 22D illustrates half-P pitch, which precesses theHeliogyro 100, 200 spin vector. The Heliogyro 100, 200 precesses, orrotates, its spin vector due to torque imparted by the sun impacting theblades 108 as they pitch along the pitch angle vs time profile 2236.Note that the pitch angle vs time sinusoidally alternates betweenpositive and negative over the duration of two rotational periods of theHeliogyro 100, 200 —meaning the blades 108 change from positivelypitched to unpitched (0°) to negatively pitched during the maneuver asthe Heliogyro 100, 200 spins twice.

Referring now to FIG. 23, a a block diagram illustrating a Heliogyro104, 204 control device in accordance with embodiments of the presentinvention is shown. The Heliogyro control device 2300 is a computingdevice including one or more processors 2304 including any processingdevices suitable for executing software applications such as Intelx86-compatible processors, embedded processors, mobile processors,and/or RISC processors.

Processor 2304 may include several devices including field-programmablegate arrays (FPGAs), memory controllers, North Bridge devices, and/orSouth Bridge devices. Although in most embodiments, processor 2304fetches application 2320 program instructions from memory 2312, itshould be understood that processor 2304 and application 2320 may beconfigured in any allowable hardware/software configuration, includingpure hardware configurations implemented in ASIC or FPGA forms. Memory2312 also includes an operating system and metadata 2316, which includesparameters and data structures used to perform the processes of thepresent application, including predetermined values and parameters invarious forms described herein.

The Heliogyro control device 2300 includes memory 2312, which mayinclude one or both of volatile and nonvolatile memory types. In someembodiments, the memory 2312 includes firmware which includes programinstructions that processor 2304 fetches and executes, including programinstructions for the operating system and software applications 2320 ofthe present application. Examples of non-volatile memory 2312 include,but are not limited to, flash memory, SD, Erasable Programmable ReadOnly Memory (EPROM), Electrically Erasable Programmable Read Only Memory(EEPROM), hard disks, and Non-Volatile Read-Only Memory (NOVRAM).Volatile memory 2312 stores various data structures and user data.Examples of volatile memory 2312 include, but are not limited to, StaticRandom Access Memory (SRAM), Dual Data Rate Random Access Memory (DDRRAM), Dual Data Rate 2 Random Access Memory (DDR2 RAM), Dual Data Rate 3Random Access Memory (DDR3 RAM), Zero Capacitor Random Access Memory(Z-RAM), Twin-Transistor Random Access Memory (TTRAM), AsynchronousRandom Access Memory (A-RAM), ETA Random Access Memory (ETA RAM), andother forms of temporary memory. Stored within memory 2312 is metadata2316 and one or more applications 2320. Metadata 2316 includes datastructures and parameters of the present application, including but notlimited to configuration parameters, predetermined values and timesettings, and information required to control a Heliogyro 100, 200.

The Heliogyro control device 2300 may include a receiver 2324 to receiveaccelerometer data 2332 from accelerometers at each blade tip 1016,blade actuator data 2336 from encoders, potentiometers, or dampers, andposition, attitude, and velocity data 2340 from a spacecraft navigationsystem 2432, an earth-based system 2420, space vehicles, satellites, orother spacecraft or any other information required to control aHeliogyro 100, 200.

The Heliogyro control device 2300 may also include a transmitter 2328 totransmit blade actuator controls 2344 to each blade actuator 124, bladedeployment actuator controls 2348 to each blade deployment actuator 708,and propulsion module controls 2352 to each propulsion module 304.

The Heliogyro control device 2300 may also include one or morecommunication transceivers 2308, which is coupled to an onboard antenna(not shown) in order to transmit or receive data from earth-basedstations 2420, space vehicles, satellites, or other spacecraft.

Referring now to FIG. 24, a diagram illustrating primary function flowfor a Heliogyro control system 2400 in accordance with embodiments ofthe present invention is shown. The Heliogyro 100, 200 includes aproprietary Guidance, Navigation, and Control (GNC) system 2400 thatautonomously provides supervised autonomy as it makes short-termtrajectory decisions and course corrections. The four major componentsof the control system 2400 are trajectory generation 2408, thrust vectorcontrol 2412, blade pitch control 2416, and state determination 2404.

Relying on state history and ground-generated updates to the a priorimission plan, a trajectory-generating algorithm 2428 calculates anoptimal trajectory to either maintain or to converge with a referencepath. Hybrid systems theory, linear covariance analysis, and numericaltargeting (shooting or collocation) may be used to calculate an updatedtrajectory. The trajectory-generating algorithm 2428 utilizes as inputsa mission plan or desired trajectory to follow Heliogyro 100, 200parameters, dynamics equations, and past Heliogyro 100, 200 behavior.Past Heliogyro behavior includes the response to previous blade pitchamounts and blade damping forces and torques. Blade 108 pitch and blade108 damping impart forces and torques to the Heliogyro 100, 200, whichin turn precesses or rotates its spin vector, changes its spin rate,changes its direction of travel, changes its velocity, or changes modalvibrations. The Heliogyro 100, 200 learns what outputs result frominputs). The trajectory-generating algorithm 2428 creates future desiredvehicle state outputs, including position, orientation, velocity, andangular velocity, and derivative of state, including acceleration andangular acceleration, vs. time.

The thrust vector controller 2412 receives current attitude, position,and velocity information 2424, as well as desired trajectory, todetermine a pattern of maneuver parameters 2452 it will send to theblade pitch controller 2456. The thrust vector controller 2412 may insome embodiments utilize an inverse lookup table that maps 2448 desiredmaneuver parameters 2452 and current Heliogyro spin axis orientation torequired thrust and moment 2444. If necessary, interim controllers canbe applied in simulation, to iteratively revise the mappings until theyare sufficient to control the Heliogyro 100, 200. New parameters aresent to the thrust vector controller 2412 at intervals of two Heliogyro100, 200 rotations. A smooth transition (through the first derivative)is applied between past pitch parameters and current parameters. Thethrust vector controller 2412 utilizes as inputs a current state, thedesired state, and a derivative of the desired state. The thrust vectorcontroller 2412 utilizes as outputs cyclic pitch amplitude, collectivepitch amplitude, half-P pitch amplitude, pitch frequency, pitch phase,and an initial time for new parameters to be applied.

The blade pitch controller 2416 generates and tracks pitch profiles forthe blade actuators 408. It receives maneuver parameters 2452 forcollective pitch angle, cyclic pitch amplitude, and phase, andtransitions from a current pitch profile to a new profile. A proprietaryrobust internal model controller, incorporating a modal model of theblade 108 assembly, pitches the blade 108 at the root while activelydamping the first twist mode.

Certain inputs to the control system—attitude, position, and velocity2424—come from a spacecraft navigation system 2432 (e.g., GEONS). Itincorporates readings such as GPS, the Tracking and Data Relay SatelliteSystem TDRSS, Doppler, and celestial navigation. In interplanetaryspace, the spacecraft navigation system 2432 typically relies primarilyon celestial navigation and Doppler. The control system 2400 thereforebenefits from highly accurate star tracking data 2436. FIG. 2illustrates an embodiment incorporating a despin mechanism. An addedbenefit of a despun section is the ability to accommodate payloads thatrequire a non-spinning platform. Also included in the control system2400 are active in-plane and out-of-plane blade dampers (not shown),located where the struts 112 attach to the spacecraft core 104, 204.

The present invention allows the overlaying of active control pitchprofiles. To limit Heliogyro 100, 200 wobble and other unwantedoscillations, techniques are incorporated from under-actuated roboticstheory to optimize damping. The techniques layer certain dampingprofiles on top of pitch controls, as well as actuate the dampers at theblade 108 roots for a system-wide optimized damping effect. Also, byspinning blades 108, the spacecraft core 104, 204 can be rapidlyprecessed. (e.g., spin two blades spaced 180 degrees apart in oppositedirections in relation to their respective local pitch axes, to quicklyturn the Heliogyro 100, 200.

Referring now to FIG. 25, a block diagram illustrating Heliogyro controlsystem model controllers in accordance with embodiments of the presentinvention is shown. The pitch actuators 124 utilize a robust internalmodel controller, with the pitch profile reference broken up into threeparts: a Half-P pitch internal model controller 2504, a cyclic pitchinternal model controller 2508, and a collective pitch internal modelcontroller 2512. A modal model of the blades 108 in twist isincorporated into the plant 2516, which allows for active damping of thepitch modes.

Referring now to FIG. 26, a diagram illustrating Heliogyro despinmechanism 212 options in accordance with embodiments of the presentinvention is shown. A Heliogyro despin mechanism 212 may include aninner core 2604 coupled to an inner gimbal, where an actuator similar toa voice coil actuator discussed with reference to the pitch actuator 124is used. One or two bearing races 2608 are positioned between the innercore 2604 and the inner gimbal to allow the inner core 2604 to rotaterelative to the inner gimbal. Slip rings may be present between thevarious axes, except between the inner core 2604 and the inner gimbal.

Referring now to FIG. 27, a diagram illustrating a side view of aHeliogyro despin mechanism 212 in accordance with embodiments of thepresent invention is shown. The despin mechanism 212 may allow power anddata to be wirelessly transferred between the spacecraft core 104, 204and an inner gimbal receiver to power and control a despin actuator. Theconfiguration can be altered to work in a CubeSat configuration.

The functional block diagrams, operational scenarios and sequences, andflow diagrams provided in the Figures are representative of exemplarysystems, environments, and methodologies for performing novel aspects ofthe disclosure. While, for purposes of simplicity of explanation,methods included herein may be in the form of a functional diagram,operational scenario or sequence, or flow diagram, and may be describedas a series of acts, it is to be understood and appreciated that themethods are not limited by the order of acts, as some acts may, inaccordance therewith, occur in a different order and/or concurrentlywith other acts from that shown and described herein. For example, thoseskilled in the art will understand and appreciate that a method couldalternatively be represented as a series of interrelated states orevents, such as in a state diagram. Moreover, not all acts illustratedin a methodology may be required for a novel embodiment.

The descriptions and figures included herein depict specific embodimentsto teach those skilled in the art how to make and use the best option.For the purpose of teaching inventive principles, some conventionalaspects have been simplified or omitted. Those skilled in the art willappreciate variations from these embodiments that fall within the scopeof the invention. Those skilled in the art will also appreciate that thefeatures described above can be combined in various ways to formmultiple embodiments. As a result, the invention is not limited to thespecific embodiments described above, but only by the claims and theirequivalents.

Finally, those skilled in the art should appreciate that they canreadily use the disclosed conception and specific embodiments as a basisfor designing or modifying other structures for carrying out the samepurposes of the present invention without departing from the spirit andscope of the invention as defined by the appended claims.

We claim:
 1. A method for controlling a spacecraft, comprising:calculating, by a control device, spacecraft position, attitude, andvelocity, the spacecraft comprising a plurality of blade actuatorscontrolling pitch for a plurality of blades, the plurality of bladesextending radially away from a spacecraft core and comprising materialconfigured to be deflected by solar pressure; receiving mission planupdates for the spacecraft; calculating an updated trajectory based onthe position, attitude, velocity, mission plan updates, and pastspacecraft behavior; generating maneuver parameters for the spacecraftfrom the updated trajectory; creating new blade pitch profiles for aplurality of blade actuators, from the maneuver parameters; sendingcontrols corresponding to the new blade pitch profiles to the pluralityof blade actuators; and transitioning from current blade pitch profilesto the new blade pitch profiles.
 2. The method for controlling aspacecraft of claim 24, wherein the mission plan updates for thespacecraft comprises instructions to either maintain or converge to areference path in space.
 3. The method for controlling a spacecraft ofclaim 24, wherein the control device receives position, attitude, andvelocity from a spacecraft navigation system, wherein the spacecraftnavigation system receives one or more of GPS readings, tracking anddata relay satellite system readings, Doppler sensor readings, startracker data, and celestial navigation readings and in responsegenerates spacecraft position, attitude, and velocity.
 4. The method forcontrolling a spacecraft of claim 24, wherein past spacecraft behaviorcomprises previous blade pitch amounts and blade damping forces andtorques.
 5. The method for controlling a spacecraft of claim 24, whereinthe control device generates maneuver parameters at intervals of twospacecraft rotations, wherein generating maneuver parameters comprising:mapping desired maneuver parameters and current spacecraft spin axisorientation to required thrust and moment through an inverse lookuptable.
 6. The method for controlling a spacecraft of claim 24, whereinthe current and new blade pitch profiles are based on collective pitchangle, cyclic pitch amplitude, and phase.
 7. The method for controllinga spacecraft of claim 24, wherein the method further comprising:overlaying active control pitch profiles on the new blade pitch profilesin order to generate damping effects for the spacecraft.